This invention relates generally to gas turbine machines and specifically, to a combustion cap assembly for a multi-nozzle, can-annular combustor.
Gas turbines generally include a compressor, one or more combustors, a fuel injection system and a turbine. Typically, the compressor pressurizes inlet air which is then turned in direction or reverse flowed to the combustors where it is used to cool the combustor and also to provide air to the combustion process. In a multi-combustor turbine, the combustors are located about the periphery of the gas turbine, and a transition duct connects the outlet end of each combustor with the inlet end of the turbine to deliver the hot products of the combustion process to the turbine.
Generally, in Dry Low NOx combustion systems utilized by the assignee, each combustor includes multiple fuel nozzles, each nozzle having a surrounding dedicated premix section or tube so that, in a premix mode, fuel is premixed with air prior to burning in the single combustion chamber. In this way, the multiple dedicated premixing sections or tubes allow thorough premixing of fuel and air prior to burning, which ultimately results in low NOx levels. See, for example, commonly owned U.S. Pat, No. 5,274,991.
More specifically, each combustor includes a generally cylindrical casing having a longitudinal axis, the casing having fore and aft sections secured to each other, and the casing as a whole secured to the turbine casing. Each combustor also includes an internal flow sleeve, and a combustion liner substantially concentrically arranged within the flow sleeve. Both the flow sleeve and combustion liner extend between the transition duct at their downstream ends, and a combustion liner cap assembly (located within an upstream portion of the combustor) at their upstream ends. The flow sleeve is attached directly to the combustor casing, while the liner supports the liner cap assembly which, in turn, is fixed to the combustor casing. The outer wall of the transition duct and at least a portion of the flow sleeve are provided with air supply holes over a substantial portion of their respective surfaces, thereby permitting compressor air to enter the radial space between the combustion liner and the flow sleeve, and to be reverse flowed to the upstream portion of the combustor.
A plurality (five in the exemplary embodiment) of diffusion/premix fuel nozzles are arranged in a circular array about the longitudinal axis of the combustor casing. These nozzles are mounted in a combustor end cover assembly which closes off the rearward end of the combustor. Inside the combustor, the fuel nozzles extend into and through the combustion liner cap assembly and, specifically, into corresponding ones of the premix tubes that are secured in the liner cap assembly. The discharge end of each nozzle terminates within a corresponding premix tube, in relatively close proximity to the downstream end of the premix tube which opens to the burning zone in the combustion liner.
Spacers between the cap""s inner body and its outer mounting flange create an annular passage for premixer air from the compressor. The premixer air travels through this annular passage, then again reverses direction within the combustor""s forward case before mixing with gaseous fuel in the inner body of the liner cap assembly, and proceeding to the reaction zone. This air flow reversal (commonly referred to as the xe2x80x9ccap turnxe2x80x9d) results in a pressure loss, which can be as high as 7% of the total combustor pressure drop. The cap turn pressure loss is a result of two effects: (1) expansion of premixer air into the forward case area after passing the cap, and (2) reversal of flow direction within the forward case to travel through the cap burner tubes.
Pressure loss in the combustor is a critical contributor to overall gas turbine performance. Any air the combustion system uses for cooling or loses to leakage is counted against the budgeted overall combustion system pressure drop.
Previous combustor designs have implemented tapered flanges on the cap assembly to allow some degree of flow expansion prior to the cap turn. However, the amount of flow expansion was relatively small, as the diffuser section was only as long as the cap mounting flange.
In this invention, the forward portion of the combustion liner cap assembly is designed as an axial diffuser to reduce the pressure drop caused by the cap turn as the premixer air passes between inner and outer bodies or components of the liner cap assembly and turns toward the fuel nozzles and the combustor chamber.
More specifically, this invention provides a combustion liner cap assembly with a conical outer body that serves to increase the cross-sectional area of the annular passage between a cylindrical inner body and the conical outer body in the direction of airflow, causing a reduction in the velocity of the premixer air as it passes through the cap assembly. These cap assembly modifications, in turn, require an enlarged forward case to accommodate the cap diffuser.
As mentioned above, the cap turn pressure loss is due to expansion of premixer air and the reversal of flow at the forward case. Since the magnitude of the pressure losses is proportional to the square of the air velocity, the reduction of air velocity caused by the axial diffuser results in a lower cap turn pressure loss. In addition, the diffuser improves flow uniformity into the premixers because the flow begins turning from the forward end of the diffuser inner cylinder rather than at the inlets to the premixer tubes. Another expected benefit of this concept is improved flame holding.
Accordingly, this invention relates to a combustion cap assembly for closing a forward end of a combustion chamber comprising a radially inner substantially cylindrical component; a radially outer substantially conical component, extending substantially along an entire length dimension of the radially inner component; and an annular airflow passage therebetween.
The invention also relates to a combustion cap assembly for rlosing a forward end of a combustion chamber comprising a radially inner substantially cylindrical component; a radially outer substantially conical component, extending substantially along an entire length dimension of the radially inner component; and an annular airflow passage therebetween; wherein the annular airflow passage increases in cross sectional area in a flow direction; and further comprising a plate supporting a plurality of premix burner tubes radially inward of said radially inner cylindrical component.
The invention also relates to a method of reducing pressure loss across a combustion liner cap assembly located on a gas turbine combustor, the cap assembly supporting a plurality of premix tubes adapted to enclose portions of a like number of nozzles, and wherein air flows in an annular passage radially outwardly of the combustor where it reverses direction to flow through the premix tubes, the method comprising adding a diffuser to the forward end of the cap assembly, the diffuser configured to increase the cross sectional area of the annular flow passage along an axial length of the cap assembly to thereby cause a reduction in velocity of the air in the annular flow passage and thereby reduce pressure loss as the air reverses direction at the forward end of the combustor.